1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to a thin wall cooled turbine rotor blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
A thin wall airfoil with near wall cooling can maintain a low metal temperature compared to thicker wall airfoil. However, thin wall airfoil cannot be cast because the liquid metal does not flow freely into the spaced formed by the ceramic cores in which forms the blade walls.
FIG. 1 shows a graph of a first stage turbine rotor blade external pressure profile used in an industrial gas turbine (IGT) engine. The forward region of the pressure side surface experiences high hot gas static pressure while the entire suction side of the airfoil is at a much lower hot gas static pressure than on the pressure side. Therefore, a higher cooling air pressured must exist on the pressure side than on the suction side of the airfoil to prevent the external hot gas from flowing into the airfoil internal cooling passages through film cooling holes.
FIG. 2 shows a prior art first stage turbine rotor blade 10 for an IGT engine with a (1+5+1) serpentine flow cooling circuit. This cooling circuit is formed as three separate sections and includes a leading edge section, a mid-chord section and a trailing edge section. the leading edge section includes a cooling air supply channel 16 connected by a row of metering and impingement holes to a leading edge impingement cavity 17 having a showerhead arrangement of film cooling holes to discharge the cooling air. The mid-chord section is cooled with a 5-pass forward flowing serpentine flow cooling circuit and includes a first leg of channel 11 located adjacent to the trailing edge section, a second leg 12, a third leg 13, a fourth leg 14 and a fifth and last leg 15 located adjacent to the leading edge cooling supply channel 16. Rows of film cooling holes are connected to the five legs and discharge onto the pressure side wall or the suction side wall or both walls. The trailing edge section is includes a cooling air supply channel 18 with a first row of metering and impingement holes 19 opening into a first diffusion cavity 20 followed by a second row of impingement holes 21 opening into a second diffusion cavity 22 followed by a row of trailing edge exit holes 23. FIG. 4 shows a flow diagram for the cooling circuit of FIG. 2.
Key design features for the prior art 5-pass serpentine flow cooling circuit used in the FIG. 2 blade include the following. Firstly, the forward flowing 5-pass serpentine is used in the airfoil mid-chord region. The cooling air flows toward and discharges into the high hot gas side pressure section of the pressure side. In order to satisfy the back flow margin criteria (no hot gas flows from external to internal of the airfoil), a high cooling air supply pressure is required and therefore induces a high leakage flow. Secondly, since the second leg and third leg of the 5-pass serpentine circuit provide film cooling air for both walls of the airfoil and in order to satisfy the back flow margin criteria for the pressure side rows of film cooling holes, the internal cavity pressure has to be approximately 10% higher than the pressure side hot gas side pressure. This results in an over-pressure of the airfoil suction side film cooling holes. Thirdly, a low aspect ration flow channels are used. this lowers the ceramic core yield (yield is the percent of non-defective cast blades) and making it difficult to install film cooling holes, a high inference due to the rotational effect on internal heat transfer coefficient, and also yields a low internal-to-hot gas side convection area ratio.
In a prior art two piece bonded blade, the airfoil pressure side piece is cast separate from the suction side piece. The two pieces are then bonded together through the use of TLP (Transient Liquid Phase) bonding. The benefits of manufacture for this blade with two piece construction is the use of a strong back ceramic core in the casting process that will allow for inspection of the internal cooling features and a measurement of the airfoil wall thickness prior to bonding the two pieces together and form the blade. However, a draw back for the two piece blade is a mismatch of the internal cold ribs, the complex trailing edge cooling features and around the airfoil edges during the bonding process.